WO2012025357A1 - Blade and corresponding fan - Google Patents

Blade and corresponding fan Download PDF

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Publication number
WO2012025357A1
WO2012025357A1 PCT/EP2011/063427 EP2011063427W WO2012025357A1 WO 2012025357 A1 WO2012025357 A1 WO 2012025357A1 EP 2011063427 W EP2011063427 W EP 2011063427W WO 2012025357 A1 WO2012025357 A1 WO 2012025357A1
Authority
WO
WIPO (PCT)
Prior art keywords
blade
tip
fan
aerofoil
tips
Prior art date
Application number
PCT/EP2011/063427
Other languages
French (fr)
Inventor
John Richard Webster
Original Assignee
Rolls-Royce Plc
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Rolls-Royce Plc filed Critical Rolls-Royce Plc
Priority to EP11739074.0A priority Critical patent/EP2609293A1/en
Priority to US13/817,587 priority patent/US20130149108A1/en
Publication of WO2012025357A1 publication Critical patent/WO2012025357A1/en

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/16—Form or construction for counteracting blade vibration
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/325—Rotors specially for elastic fluids for axial flow pumps for axial flow fans
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—Blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/38—Blades
    • F04D29/384—Blades characterised by form
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
    • F04D29/668—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps damping or preventing mechanical vibrations
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/183—Two-dimensional patterned zigzag
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/70—Shape

Definitions

  • the present invention relates to a blade, for example to a fan blade for a turbofan gas turbine engine.
  • Fan flutter and other vibration continues to be a significant issue.
  • the traditional route to reduce this is to avoid running range/blade or fan set modes, but this is particularly difficult at take off.
  • Alternative methods include re-camber and increased blade chord.
  • Turbofan clapperless fan blades may suffer from vibration where aerodynamic forces lead to excitation of a fan blades natural modes of vibration, e.g. second flap mode, away from coincidence with the harmonics of a fan blades rotational speed, i.e. a non integral vibration.
  • Avoidance of flutter mode coincidences restricts running range, recamber reduces efficiency, and additional chord increases weight.
  • the present invention seeks to provide a novel blade, which at least reduces the above problem.
  • the present invention provides a blade comprising a root portion and an aerofoil portion, wherein the aerofoil portion has a tip remote from the root portion, and a leading edge and a trailing edge, and wherein the tip of the aerofoil portion has a set-back portion extending from the leading edge or the trailing edge of the aerofoil portion part way towards the respective other edge and set back from the remainder of the tip of the aerofoil portion towards the root portion.
  • the set-back portion in the tip is serrated, more preferably with serration slots shaped and not aligned with the circumferential direction of motion of the tip when the blade is rotating in use.
  • the serration slots may be approximately perpendicular to the surface of the tip which is flow-washed when the blade is rotating in use in a fan.
  • the serration slots are at most 2mm deep.
  • the blade is a fan blade.
  • the present invention also provides a fan having a plurality of blades in accordance with the blade invention as set out above and a fan casing around the tips of the blades, wherein as a result of the set-back portions in the tips the tip clearance area between tips and fan casing is changed by at least 1 % of fan area as compared with a case in which the set-back portions in the tips were omitted.
  • the present invention also provides an engine, for example a turbofan gas turbine engine, having a blade or a fan in accordance with the blade invention or fan invention as set out above.
  • an engine for example a turbofan gas turbine engine, having a blade or a fan in accordance with the blade invention or fan invention as set out above.
  • embodiments of the present invention can provide for blade vibration damping by utilising passive modulation of blade tip clearance.
  • Embodiments of the present invention can provide for extended blade life due to reduction in high cycle fatigue, reduced blade generated noise due to blade damping, reduced blade tip generated noise due to disrupted over tip vortex. With embodiments of the present invention problems of reduced fan efficiency and/or increased weight can be at least mitigated.
  • Tip clearance modulation in accordance with embodiments of the present inventions can have a significant effect on blade vibration, for example in fans and/or compressors.
  • Figure 1 shows a turbofan gas turbine engine having a fan blade to which the present invention can be applied.
  • Figure 2 shows a fan blade to which the present invention can be applied.
  • Figure 3 schematically illustrates a simplified tip modulation scenario, for assistance in understanding the present invention.
  • Figure 4 schematically illustrates tip opening on a twisted fan blade for assistance in understanding the present invention.
  • FIGS 5 and 6 schematically illustrate blades in accordance with embodiments of the invention.
  • Figures 7 and 8 show graphs relating to the present invention.
  • a turbofan gas turbine engine 10 as shown in Fig. 1 , comprises in flow series an inlet 12, a fan section 14, a compressor section 16, a combustion section 18, a turbine section 20 and an exhaust 22.
  • the fan section 14 comprises a fan rotor 24 carrying a plurality of circumferentially spaced radially outwardly extending fan blades 26.
  • the fan blades 26 are arranged in a bypass duct 28 defined by a fan casing 30, which surrounds the fan rotor 24 and fan blades 26.
  • the fan casing 30 is secured to a core engine casing 34 by a plurality of circumferentially spaced radially extending fan outlet guide vanes 32.
  • the fan rotor 24 and fan blades 26 are arranged to be driven by a turbine (not shown) in the turbine section 20 via a shaft (not shown).
  • the compressor section 16 comprises one or more compressors (not shown) arranged to be driven by one or more turbines (not shown) in the turbine section 20 via respective shafts (not shown).
  • the fan blade 26 comprises a root portion 36 and an aerofoil portion 38.
  • the root portion 36 is arranged to locate in a slot 40 in the rim 42 of the fan rotor 24, and for example the root portion 36 may be dovetail shape, or fir-tree shape, in cross-section and hence the corresponding slot 40 in the rim 42 of the fan rotor 24 is the same shape.
  • the aerofoil portion 38 has a leading edge 44, a trailing edge 46 and a tip 48 remote from the root portion 36 and the fan rotor 24.
  • a concave pressure surface 50 extends from the leading edge 44 to the trailing edge 46 and a convex suction surface 52 extends from the leading edge 44 to the trailing edge 46.
  • the inventor has had the insight that aerodynamic disturbances caused by vibration of the blades 26 could excite appropriate modes in the casing 30 that would in turn modulate the tip clearance. It is suspected that changes in tip clearance cause a modulation in the energy loss due to tip leakage and hence a modulation in the aerodynamic loading, particularly around the tip 48. This loading modulation can provide a vibration excitation. Dependent on modal coincidences, mode strengths and exact phasing, the mechanism can provide strong excitation or damping.
  • the inventor has further had the insight that an asymmetric tip blade can provide an effect affording correct modes and frequencies, which can be relatively insensitive to exact conditions and is easier to incorporate into new or existing designs.
  • a modulation in this energy loss can provide vibration forcing/damping.
  • Fig. 3 which schematically illustrates tip modulation considering a blade as a simple flat plate, which operates close to a flat plate (casing) - a flap mode will provide a tip clearance modulation. This modulation opens the gap at the maximum displacement on each half-vibration cycle, so that the modulation occurs at twice the vibration frequency.
  • Fig. 4 schematically illustrates tip opening on a twisted fan blade.
  • Fig. 5 illustrates a blade in accordance with an embodiment of the invention, in this case a blade configured at the tip (uppermost in the Figure) with a set-back portion (54) to give increased (tip) clearance towards trailing edge - other embodiments may reverse the profile (e.g. to give increased (tip) clearance towards leading edge).
  • the set-back portion may for example be dimensioned to increase tip clearance area (compared to a tip without set-back portion) vis-a-vis the casing (not shown) equivalent to 1 % of fan area.
  • the blade comprises a root portion 36 and an aerofoil portion 38, the aerofoil portion 38 having a tip 48 remote from the root portion 36, and a leading edge 44 and a trailing edge 48.
  • the tip 48 of the aerofoil portion 38 has a set-back portion 54 extending from the leading edge 44 (or the trailing edge 48 in the case of a reversed profile) of the aerofoil portion 38 part way towards the respective other edge 48; 44 and set back from the remainder of the tip 48 of the aerofoil portion 38 towards the root portion 36.
  • the set-back portion 54 of the tip 48 is serrated - see Fig. 6 (a blade with serrated tip can provide increased clearance and ability to cut lining - other embodiments may again reverse the profile, e.g. to give increased (tip) clearance towards leading edge)- so that it would still cut the lining to the same depth, but give an increased over tip leakage equivalent to an increased clearance. Serrations a few mm deep, e.g. from 4mm deep to 3mm deep, or to as little as 2mm deep would be adequate.
  • the inventor has realized that the aerodynamic effect of dynamic changes in tip clearance may in some cases be initially detrimental, but if the serration slots are shaped and not aligned with the circumferential direction of motion of the tip when the blade is rotating in use an efficiency benefit can be reestablished. It is important to know the efficiency of the control effect and the phase lag between the clearance modulation and the blade forcing. As described above, a 180° phase change can be obtained, so some benefit is achieved even if an exact phase match between excitation and required damping is not precisely known.
  • a steady state tip clearance area change equivalent to 1 % of fan area gives a significant efficiency change.
  • a ⁇ 0.5mm tip clearance change might produce a change in output power of 170kW.
  • first flap a typical blade has a blade energy in the order of 60J at a modest amplitude.
  • a Q factor of around 60 must be achieved to give an acceptable level of damping.
  • the present invention is for example applicable to clapperless fan blades which lead to excitation of other natural modes of vibration, e.g. first flap mode, third flap mode, first torsion mode, second torsion mode or combinations thereof or any of the first ten fundamental vibration modes.
  • the present invention is applicable to metal fan blades and hybrid structured fan blades e.g. composite fan blades. In the case of some designs of hybrid structured fan blades there may be other natural modes of vibration that are not easy to describe using first flap mode, second flap mode, third flap mode, first torsion mode or second torsion mode because the complex structure of these hybrid structured fan blades may distort such mode shapes out of recognition.
  • the present invention is however also applicable to other fan or turbine applications or turbomachinery blades, including e. g. fans in ventilation subsystems or automotive applications, centrifugal compressors etc.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A blade comprising a root portion (36) and an aerofoil portion(38), wherein the aerofoil portion (38) has a tip (48) remote from the root portion (36), and a leading edge (44) and a trailing edge (48), and wherein the tip (48) of the aerofoil portion (38) has a set-back portion (54) extending from the leading edge (44) or the trailing edge (48) of the aerofoil portion (38) part way towards the respective other edge (48; 44) and set back from the remainder of the tip (48) of the aerofoil portion (38) towards the root portion (36).

Description

BLADE AND CORRESPONDING FAN
The present invention relates to a blade, for example to a fan blade for a turbofan gas turbine engine.
Fan flutter and other vibration continues to be a significant issue. The traditional route to reduce this is to avoid running range/blade or fan set modes, but this is particularly difficult at take off. Alternative methods include re-camber and increased blade chord.
Turbofan clapperless fan blades may suffer from vibration where aerodynamic forces lead to excitation of a fan blades natural modes of vibration, e.g. second flap mode, away from coincidence with the harmonics of a fan blades rotational speed, i.e. a non integral vibration.
Avoidance of flutter mode coincidences restricts running range, recamber reduces efficiency, and additional chord increases weight.
Accordingly the present invention seeks to provide a novel blade, which at least reduces the above problem.
Accordingly the present invention provides a blade comprising a root portion and an aerofoil portion, wherein the aerofoil portion has a tip remote from the root portion, and a leading edge and a trailing edge, and wherein the tip of the aerofoil portion has a set-back portion extending from the leading edge or the trailing edge of the aerofoil portion part way towards the respective other edge and set back from the remainder of the tip of the aerofoil portion towards the root portion.
Preferably the set-back portion in the tip is serrated, more preferably with serration slots shaped and not aligned with the circumferential direction of motion of the tip when the blade is rotating in use.
The serration slots may be approximately perpendicular to the surface of the tip which is flow-washed when the blade is rotating in use in a fan.
Preferably the serration slots are at most 2mm deep.
Preferably the blade is a fan blade.
The present invention also provides a fan having a plurality of blades in accordance with the blade invention as set out above and a fan casing around the tips of the blades, wherein as a result of the set-back portions in the tips the tip clearance area between tips and fan casing is changed by at least 1 % of fan area as compared with a case in which the set-back portions in the tips were omitted.
The present invention also provides an engine, for example a turbofan gas turbine engine, having a blade or a fan in accordance with the blade invention or fan invention as set out above.
In summary, embodiments of the present invention can provide for blade vibration damping by utilising passive modulation of blade tip clearance. Embodiments of the present invention can provide for extended blade life due to reduction in high cycle fatigue, reduced blade generated noise due to blade damping, reduced blade tip generated noise due to disrupted over tip vortex. With embodiments of the present invention problems of reduced fan efficiency and/or increased weight can be at least mitigated. Tip clearance modulation in accordance with embodiments of the present inventions can have a significant effect on blade vibration, for example in fans and/or compressors.
Exemplary embodiments of the present invention will be more fully described by way of example with reference to the accompanying drawings in which:
Figure 1 shows a turbofan gas turbine engine having a fan blade to which the present invention can be applied.
Figure 2 shows a fan blade to which the present invention can be applied.
Figure 3 schematically illustrates a simplified tip modulation scenario, for assistance in understanding the present invention.
Figure 4 schematically illustrates tip opening on a twisted fan blade for assistance in understanding the present invention.
Figures 5 and 6 schematically illustrate blades in accordance with embodiments of the invention.
Figures 7 and 8 show graphs relating to the present invention.
A turbofan gas turbine engine 10, as shown in Fig. 1 , comprises in flow series an inlet 12, a fan section 14, a compressor section 16, a combustion section 18, a turbine section 20 and an exhaust 22. The fan section 14 comprises a fan rotor 24 carrying a plurality of circumferentially spaced radially outwardly extending fan blades 26. The fan blades 26 are arranged in a bypass duct 28 defined by a fan casing 30, which surrounds the fan rotor 24 and fan blades 26. The fan casing 30 is secured to a core engine casing 34 by a plurality of circumferentially spaced radially extending fan outlet guide vanes 32. The fan rotor 24 and fan blades 26 are arranged to be driven by a turbine (not shown) in the turbine section 20 via a shaft (not shown). The compressor section 16 comprises one or more compressors (not shown) arranged to be driven by one or more turbines (not shown) in the turbine section 20 via respective shafts (not shown).
An exemplary fan blade 26 to which the present invention can be applied is shown more clearly in Fig. 2. The fan blade 26 comprises a root portion 36 and an aerofoil portion 38. The root portion 36 is arranged to locate in a slot 40 in the rim 42 of the fan rotor 24, and for example the root portion 36 may be dovetail shape, or fir-tree shape, in cross-section and hence the corresponding slot 40 in the rim 42 of the fan rotor 24 is the same shape. The aerofoil portion 38 has a leading edge 44, a trailing edge 46 and a tip 48 remote from the root portion 36 and the fan rotor 24. A concave pressure surface 50 extends from the leading edge 44 to the trailing edge 46 and a convex suction surface 52 extends from the leading edge 44 to the trailing edge 46.
The inventor has had the insight that aerodynamic disturbances caused by vibration of the blades 26 could excite appropriate modes in the casing 30 that would in turn modulate the tip clearance. It is suspected that changes in tip clearance cause a modulation in the energy loss due to tip leakage and hence a modulation in the aerodynamic loading, particularly around the tip 48. This loading modulation can provide a vibration excitation. Dependent on modal coincidences, mode strengths and exact phasing, the mechanism can provide strong excitation or damping.
The inventor has further had the insight that an asymmetric tip blade can provide an effect affording correct modes and frequencies, which can be relatively insensitive to exact conditions and is easier to incorporate into new or existing designs.
The inventor has appreciated that small changes in tip clearance can cause major performance penalties i.e. energy loss. This energy loss will be manifested as a reduction of the blade loading around the tip.
Expressed very briefly the inventor has realized that a modulation in this energy loss can provide vibration forcing/damping. As a simplified illustration - see Fig. 3, which schematically illustrates tip modulation considering a blade as a simple flat plate, which operates close to a flat plate (casing) - a flap mode will provide a tip clearance modulation. This modulation opens the gap at the maximum displacement on each half-vibration cycle, so that the modulation occurs at twice the vibration frequency.
Since this is frequency doubled, it can have no effect on the blade vibration in the flap mode. However, the inventor has had the insight that if some asymmetry is introduced the modulation can be made to occur only once per cycle. This configuration now has the potential to provide an aerodynamic forcing which is at the same frequency as the blade vibration. The phase of this forcing can be changed by 180° to provide damping.
The real situation is more complex than is illustrated in Fig. 3, involving a curved casing and in the case of the fan blade, high levels of blade twist, which gives significant modification to the tip motion. The effect will increase towards the leading and trailing edges. In the case of a twisted blade, the motion is not perpendicular to the tip aerofoil with modulation once per cycle. Fig. 4 schematically illustrates tip opening on a twisted fan blade.
With a simple model, as the inventor has realized, the effect from the leading and trailing edges would however be equal and opposite so would cancel each other out. The inventor has further appreciated that asymmetry in geometry or local aerodynamic loading could lead to an out of balance effect that will result in blade forcing and suspects that this is likely to occur in existing designs and may be the root of some vibration problems. However the inventor has had the further insight that the effect could be enhanced by deliberately increasing the clearance towards the leading or trailing edge and that this would reduce the effect in that region, leaving the other edge to dominate and provide a useful effect.
Fig. 5 illustrates a blade in accordance with an embodiment of the invention, in this case a blade configured at the tip (uppermost in the Figure) with a set-back portion (54) to give increased (tip) clearance towards trailing edge - other embodiments may reverse the profile (e.g. to give increased (tip) clearance towards leading edge). The set-back portion may for example be dimensioned to increase tip clearance area (compared to a tip without set-back portion) vis-a-vis the casing (not shown) equivalent to 1 % of fan area. Thus, the blade comprises a root portion 36 and an aerofoil portion 38, the aerofoil portion 38 having a tip 48 remote from the root portion 36, and a leading edge 44 and a trailing edge 48. The tip 48 of the aerofoil portion 38 has a set-back portion 54 extending from the leading edge 44 (or the trailing edge 48 in the case of a reversed profile) of the aerofoil portion 38 part way towards the respective other edge 48; 44 and set back from the remainder of the tip 48 of the aerofoil portion 38 towards the root portion 36.
In another embodiment, the set-back portion 54 of the tip 48 is serrated - see Fig. 6 (a blade with serrated tip can provide increased clearance and ability to cut lining - other embodiments may again reverse the profile, e.g. to give increased (tip) clearance towards leading edge)- so that it would still cut the lining to the same depth, but give an increased over tip leakage equivalent to an increased clearance. Serrations a few mm deep, e.g. from 4mm deep to 3mm deep, or to as little as 2mm deep would be adequate.
The inventor has realized that the aerodynamic effect of dynamic changes in tip clearance may in some cases be initially detrimental, but if the serration slots are shaped and not aligned with the circumferential direction of motion of the tip when the blade is rotating in use an efficiency benefit can be reestablished. It is important to know the efficiency of the control effect and the phase lag between the clearance modulation and the blade forcing. As described above, a 180° phase change can be obtained, so some benefit is achieved even if an exact phase match between excitation and required damping is not precisely known.
In an example, for a large turbofan engine, a steady state tip clearance area change equivalent to 1 % of fan area gives a significant efficiency change. For example for a 2.5m fan using 60MW of power, a ± 0.5mm tip clearance change might produce a change in output power of 170kW. In first flap, a typical blade has a blade energy in the order of 60J at a modest amplitude. A Q factor of around 60 must be achieved to give an acceptable level of damping.
From the basic equation
2 x KE x p (where KE is kinetic energy/blade energy and p is pi) the loss per set of blades must be in the order of 7kW.
Based on these approximations, there is needed to achieve a damping effect of 7kW from a potential energy input of 170kW i.e. 4% efficiency.
Since this would require an increase in average tip clearance of only
0.5mm, it would result in a modest performance loss which might be gained by redesign of other blade features.
If greater than 4% efficiency could be achieved on the basic mechanism, this performance loss can be reduced.
In the graphs of Figs. 7 and 8, data on tip clearance modulation is indicated. The data in the graph of Fig. 7 presents trailing edge radial movement for several modes, the flap mode showing a clear period of approximately 7 ms. The clearance closure is at blade frequency and shows no evidence of frequency doubling.
Taking the blade and casing geometry into account, the motion can be plotted relative to the casing as shown in the graph of Fig. 8. This shows a small but clear change in radial (Z) motion relative to the mean radius as the blade moves tangentially. The phasing of the leading and trailing edges being clearly opposite.
The present invention is for example applicable to clapperless fan blades which lead to excitation of other natural modes of vibration, e.g. first flap mode, third flap mode, first torsion mode, second torsion mode or combinations thereof or any of the first ten fundamental vibration modes. The present invention is applicable to metal fan blades and hybrid structured fan blades e.g. composite fan blades. In the case of some designs of hybrid structured fan blades there may be other natural modes of vibration that are not easy to describe using first flap mode, second flap mode, third flap mode, first torsion mode or second torsion mode because the complex structure of these hybrid structured fan blades may distort such mode shapes out of recognition.
The present invention is however also applicable to other fan or turbine applications or turbomachinery blades, including e. g. fans in ventilation subsystems or automotive applications, centrifugal compressors etc.

Claims

Claims
1 . A blade comprising a root portion (36) and an aerofoil portion (38), wherein the aerofoil portion (38) has a tip (48) remote from the root portion (36), and a leading edge (44) and a trailing edge (48), and wherein the tip (48) of the aerofoil portion (38) has a set-back portion (54) extending from the leading edge (44) or the trailing edge (48) of the aerofoil portion (38) part way towards the respective other edge (48; 44) and set back from the remainder of the tip (48) of the aerofoil portion (38) towards the root portion (36).
2. A blade as claimed in claim 1 , wherein the set-back portion (54) in the tip (48) is serrated.
3. A blade as claimed in claim 2, wherein the serrated set-back portion (54) has shaped serration slots which extend not aligned with the circumferential direction of motion of the tip (48) when the blade is rotating in use in a fan.
4. A blade as claimed in claim 3, wherein the serration slots are approximately perpendicular to the surface of the tip (48) which is flow-washed when the blade is rotating in use in a fan.
5. A blade as claimed in claim 3 or 4, wherein the serration slots are 2mm deep.
6. A fan having a plurality of blades as claimed in any preceding claim and a fan casing (30) around the tips (48) of the blades (26), wherein as a result of the setback portions (54) in the tips (48) the tip clearance area between tips and fan casing (30) is changed by at least 1 % of fan area as compared with a case in which the set-back portions (54) in the tips (48) were omitted.
7. A blade substantially as described in this specification, with reference to and as shown in Figures 3 to 8 of the accompanying drawings.
PCT/EP2011/063427 2010-08-23 2011-08-04 Blade and corresponding fan WO2012025357A1 (en)

Priority Applications (2)

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GB1014019.2 2010-08-23

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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2014049239A1 (en) 2012-09-25 2014-04-03 Snecma Turbomachine casing and impeller
US9568009B2 (en) 2013-03-11 2017-02-14 Rolls-Royce Corporation Gas turbine engine flow path geometry

Families Citing this family (30)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9102397B2 (en) * 2011-12-20 2015-08-11 General Electric Company Airfoils including tip profile for noise reduction and method for fabricating same
CN104314868B (en) * 2012-04-10 2017-07-14 夏普株式会社 Propeller type fan, fluid delivery system, electric fan and molding die
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CN105658038B (en) * 2016-03-18 2020-12-18 联想(北京)有限公司 Heat dissipation device and electronic equipment
US20180112542A1 (en) * 2016-10-24 2018-04-26 Pratt & Whitney Canada Corp. Gas turbine engine rotor
CN107762973B (en) * 2017-10-20 2020-06-16 哈尔滨工程大学 Compressor corner region stability-expanding blade and trailing edge groove forming method thereof
US20200224669A1 (en) * 2019-01-11 2020-07-16 Dyna Rechi Co., Ltd. Fan blade structure

Citations (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB753652A (en) * 1951-05-25 1956-07-25 Vladimir Henry Pavlecka A method of compressing a fluid
DE1057137B (en) * 1958-03-07 1959-05-14 Maschf Augsburg Nuernberg Ag Blade gap seal on centrifugal machines with impellers without a cover band or cover disk
DE2034890A1 (en) * 1969-07-21 1971-02-04 Rolls Royce Ltd Derby, Derbyshire (Großbritannien) Blade for axial flow machines
GB2034435A (en) * 1978-10-24 1980-06-04 Gerry U Fluid rotary power conversion means
US4497613A (en) * 1983-01-26 1985-02-05 General Electric Company Tapered core exit for gas turbine bucket
US4738586A (en) * 1985-03-11 1988-04-19 United Technologies Corporation Compressor blade tip seal
EP0675290A2 (en) * 1994-03-28 1995-10-04 Research Institute Of Advanced Material Gas-Generator, Ltd. Axial flow compressor
JPH0913904A (en) * 1995-06-27 1997-01-14 Ishikawajima Harima Heavy Ind Co Ltd Ceramic turbine moving blade
US6059532A (en) * 1997-10-24 2000-05-09 Alliedsignal Inc. Axial flow turbo-machine fan blade having shifted tip center of gravity axis
US6338609B1 (en) * 2000-02-18 2002-01-15 General Electric Company Convex compressor casing
US20030059309A1 (en) * 2001-09-26 2003-03-27 Szucs Peter Nicholas Methods and apparatus for improving engine operation
US20040028526A1 (en) * 2002-08-09 2004-02-12 Honda Giken Kogyo Kabushiki Kaisha Axial flow compressor
US20050106030A1 (en) * 2003-11-08 2005-05-19 Rene Bachofner Compressor rotor blade
US20070020101A1 (en) * 2005-07-22 2007-01-25 United Technologies Corporation Fan rotor design for coincidence avoidance
EP1918525A2 (en) * 2006-10-16 2008-05-07 United Technologies Corporation Gas turbine engine seal and corresponding passive sealing clearance control system
WO2008082397A1 (en) * 2006-12-29 2008-07-10 Carrier Corporation Reduced tip clearance losses in axial flow fans
EP1985805A1 (en) * 2007-04-26 2008-10-29 Siemens Aktiengesellschaft Rotary machine
EP2141328A1 (en) * 2008-07-03 2010-01-06 Siemens Aktiengesellschaft Sealing system between a shroud segment and a rotor blade tip and manufacturing method for such a segment
EP2146053A1 (en) * 2008-07-17 2010-01-20 Siemens Aktiengesellschaft Axial turbomachine with low tip leakage losses
WO2011026468A2 (en) * 2009-09-04 2011-03-10 Mtu Aero Engines Gmbh Turbomachine, and method for producing a structured abradable coating

Family Cites Families (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1317707A (en) * 1919-10-07 Inghouse electric
US1862827A (en) * 1930-01-22 1932-06-14 Parsons Steam turbine
US2976352A (en) * 1957-11-14 1961-03-21 Torrington Mfg Co Blower unit
US4274806A (en) * 1979-06-18 1981-06-23 General Electric Company Staircase blade tip
US4692098A (en) * 1981-08-31 1987-09-08 General Motors Corporation Airfoil for high efficiency/high lift fan
FR2615254A1 (en) * 1987-05-13 1988-11-18 Snecma MOBILE BLOWER BLADE COMPRISING AN END END
US5215441A (en) * 1991-11-07 1993-06-01 Carrier Corporation Air conditioner with condensate slinging fan
US5452575A (en) * 1993-09-07 1995-09-26 General Electric Company Aircraft gas turbine engine thrust mount
DE19546008A1 (en) * 1995-12-09 1997-06-12 Abb Patent Gmbh Turbine blade, which is intended for use in the wet steam area of pre-output and output stages of turbines
JP3916723B2 (en) * 1997-05-15 2007-05-23 富士重工業株式会社 Rotor blade of rotorcraft
US6027306A (en) * 1997-06-23 2000-02-22 General Electric Company Turbine blade tip flow discouragers
US5997251A (en) * 1997-11-17 1999-12-07 General Electric Company Ribbed turbine blade tip
US6223524B1 (en) * 1998-01-23 2001-05-01 Diversitech, Inc. Shrouds for gas turbine engines and methods for making the same
US6139019A (en) * 1999-03-24 2000-10-31 General Electric Company Seal assembly and rotary machine containing such seal
US6428278B1 (en) * 2000-12-04 2002-08-06 United Technologies Corporation Mistuned rotor blade array for passive flutter control
US6761539B2 (en) * 2002-07-24 2004-07-13 Ventilatoren Sirocco Howden B.V. Rotor blade with a reduced tip
US6902377B2 (en) * 2003-04-21 2005-06-07 Intel Corporation High performance axial fan
GB0513377D0 (en) * 2005-06-30 2005-08-03 Rolls Royce Plc A blade
GB2461502B (en) * 2008-06-30 2010-05-19 Rolls Royce Plc An aerofoil
US8662834B2 (en) * 2009-06-30 2014-03-04 General Electric Company Method for reducing tip rub loading
US8556575B2 (en) * 2010-03-26 2013-10-15 United Technologies Corporation Blade outer seal for a gas turbine engine
GB201006449D0 (en) * 2010-04-19 2010-06-02 Rolls Royce Plc Blades

Patent Citations (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB753652A (en) * 1951-05-25 1956-07-25 Vladimir Henry Pavlecka A method of compressing a fluid
DE1057137B (en) * 1958-03-07 1959-05-14 Maschf Augsburg Nuernberg Ag Blade gap seal on centrifugal machines with impellers without a cover band or cover disk
DE2034890A1 (en) * 1969-07-21 1971-02-04 Rolls Royce Ltd Derby, Derbyshire (Großbritannien) Blade for axial flow machines
GB2034435A (en) * 1978-10-24 1980-06-04 Gerry U Fluid rotary power conversion means
US4497613A (en) * 1983-01-26 1985-02-05 General Electric Company Tapered core exit for gas turbine bucket
US4738586A (en) * 1985-03-11 1988-04-19 United Technologies Corporation Compressor blade tip seal
EP0675290A2 (en) * 1994-03-28 1995-10-04 Research Institute Of Advanced Material Gas-Generator, Ltd. Axial flow compressor
JPH0913904A (en) * 1995-06-27 1997-01-14 Ishikawajima Harima Heavy Ind Co Ltd Ceramic turbine moving blade
US6059532A (en) * 1997-10-24 2000-05-09 Alliedsignal Inc. Axial flow turbo-machine fan blade having shifted tip center of gravity axis
US6338609B1 (en) * 2000-02-18 2002-01-15 General Electric Company Convex compressor casing
US20030059309A1 (en) * 2001-09-26 2003-03-27 Szucs Peter Nicholas Methods and apparatus for improving engine operation
US20040028526A1 (en) * 2002-08-09 2004-02-12 Honda Giken Kogyo Kabushiki Kaisha Axial flow compressor
US20050106030A1 (en) * 2003-11-08 2005-05-19 Rene Bachofner Compressor rotor blade
US20070020101A1 (en) * 2005-07-22 2007-01-25 United Technologies Corporation Fan rotor design for coincidence avoidance
EP1918525A2 (en) * 2006-10-16 2008-05-07 United Technologies Corporation Gas turbine engine seal and corresponding passive sealing clearance control system
WO2008082397A1 (en) * 2006-12-29 2008-07-10 Carrier Corporation Reduced tip clearance losses in axial flow fans
EP1985805A1 (en) * 2007-04-26 2008-10-29 Siemens Aktiengesellschaft Rotary machine
EP2141328A1 (en) * 2008-07-03 2010-01-06 Siemens Aktiengesellschaft Sealing system between a shroud segment and a rotor blade tip and manufacturing method for such a segment
EP2146053A1 (en) * 2008-07-17 2010-01-20 Siemens Aktiengesellschaft Axial turbomachine with low tip leakage losses
WO2011026468A2 (en) * 2009-09-04 2011-03-10 Mtu Aero Engines Gmbh Turbomachine, and method for producing a structured abradable coating

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2014049239A1 (en) 2012-09-25 2014-04-03 Snecma Turbomachine casing and impeller
US9982554B2 (en) 2012-09-25 2018-05-29 Snecma Turbine engine casing and rotor wheel
US9568009B2 (en) 2013-03-11 2017-02-14 Rolls-Royce Corporation Gas turbine engine flow path geometry

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US20130149108A1 (en) 2013-06-13
EP2609293A1 (en) 2013-07-03
GB201014019D0 (en) 2010-10-06

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